43results about How to "Guaranteed aerodynamic shape" patented technology

The invention discloses an airship with an inflating skeleton expanding in air, which comprises an airship body 1, a tail wing 2, a tail thruster 3, a connecting element 7, a payload bay 8, a solar cell panel 9, a nose cone 10, a first auxiliary rope 11, an automatic inflating device 12, a second auxiliary rope 13 and two high-altitude balloons 14. The airship is characterized in that a dimensional element of the airship body 1 is skin 4, a longitudinal element is a rigid skeleton 5, and an annular element is an inflating skeleton 6; the inflating skeleton 6 is connected with the rigid skeleton 5 by the connecting element; and the skin 4 is fixedly connected on the outer surface of a grid formed by the rigid skeleton 5 and the inflating skeleton 6. The invention overcomes the defect that a soft bag in a folding state and a connecting part of a pod and other rigid structures are easily torn by strong wind and vortex in the lifting process.

The invention discloses a full-active cooling hypersonic flight vehicle which comprises a transpiration cooling end, a transpiration cooling wing front edge, a transpiration cooling rudder front edge,a regenerative cooling rudder surface, a regenerative cooling windward side and a flight vehicle leeside, wherein porous structures are formed in the transpiration cooling end, the transpiration cooling wing front edge and the transpiration cooling rudder front edge, and the interiors of the porous structures are filled with liquid cooling media; the liquid cooling media can be supplied outwardsfrom the porous structures, flow through the porous structures to perform heat exchange, are heated and then flow out of the outer surface of the end in a gaseous form; and the regenerative cooling rudder surface and the regenerative cooling windward side are respectively used for cooling the rudder surface and the windward side of the flight vehicle in a regenerative cooling manner, and the liquid cooling media carried inside the flight vehicle are used for blocking, cooling and taking away pneumatic heating volume of the flight vehicle transferred into the flight vehicle in the hypersonic flight process, so that thermal protection requirements of the flight vehicle on high-speed, long-term and long-distance flight in the atmosphere are met.

The embodiment of the invention discloses a precision casting forming technology for a high-strength aluminum alloy of a large thin-walled wing body fusion special-shaped cabin shell. The technology comprises the following steps of casting a ZL205A high-strength aluminum alloy large thin-walled mesh rib wing body fusion cabin shell through an integral precision sand casting molding process; performing heat treatment on a large thin-walled aluminum alloy special-shaped cabin shell and controlling a deformation amount by adjusting parameters in a heat treatment process to form a high-performancecasting after the heat treatment; and performing numerical control machining on the casting which is subjected to the heat treatment, and controlling the deformation amount through integral surrounding layered cutting cycle machining, three-dimensional photographing scanning, datum transforming and the like in a numerical control machining process, so that precision machining deformation controlof a large thin-walled weakly rigid special-shaped complex cabin product is completed. Through the technology, series of key technical problems that the conventional aluminum alloy casting forming process cannot realize ZL205A high-strength aluminum alloy precision casting and the heat treatment of a large thin-walled weakly rigid mesh rib special-shaped cabin body, and the machining deformation is difficult to control are solved.

The invention belongs to the technical field of machining and manufacturing, and discloses a manufacturing method for an ultra-long thin-wall aluminum alloy sandwich grid rib double-skin integral tailfin. The manufacturing method includes the steps that a ZL205AA high-strength aluminum alloy is used for pouring and molding the integral tail fin; heat treatment is carried out on the integral tailfin; and numerical control machining is carried out on the integral tail fin after heat treatment. According to the manufacturing method for the ultra-long thin-wall aluminum alloy sandwich grid rib double-skin integral tail fin, the overall quality of precision molding, machining and manufacturing can be improved, and the manufacturing cycle is shortened while the manufacturing cost is lowered.

The invention discloses a water-air dual-use unmanned aerial vehicle. The water-air dual-use unmanned aerial vehicle includes a fuselage, wings, tail wings, an air propeller, a folding and unfolding mechanism, booster mechanisms, a control surface driving mechanism, a shedding mechanism and navigation propellers; the fuselage is a main mounting body and connected the wings, the tail wings, the airpropeller, the folding and unfolding mechanism, the shedding mechanism and the navigation propellers; the folding mechanism is used for folding and unfolding the wings; the booster mechanisms providethe energy for underwater emission of the unmanned aerial vehicle; the control surface driving mechanism is used for the control of each control surface; the shedding mechanism is used for the redundant booting mechanisms during the unmanned aerial vehicle flight; and the navigation propellers are used for underwater navigation. According to the water-air dual-use unmanned aerial vehicle, the folding and unfolding mechanism is used for folding the wings backwards, the resistance of the underwater navigation is reduced, an air cylinder is adopted to boost, the launching force is provided, thecontrol surface driving mechanism with small occupied space volume is adopted to ensure water-proof sealing, the shedding mechanism is adopted to shed the air cylinder, and the unmanned aerial vehicleweight and flight resistance are reduced.

The invention belongs to the technical field of guided missile structure design, and relates to a longitudinal unfolding mechanism of a missile folding rudder surface. The mechanism comprises a folding rudder surface and a fixed rudder surface; the leading edge of the folding rudder surface stretches out of a rudder tiller, the rudder tiller is provided with a rotating shaft hole, and the rotatingshaft hole is provided with a bearing; the rotating shaft hole is matched with a bearing outer ring, an inner ring of the bearing cooperates with a mandrel; the folding rudder surface is connected with the fixed rudder surface through the mandrel and the bearing, and the folding rudder surface can be folded forward to the surface of a stern room around the mandrel; when the folding rudder surfaceis in a folded state, a pin puller arranged in the stern room is inserted into a square-shaped groove in the front edge of the folding rudder surface, and the folding rudder surface is locked; a tension spring is connected between the folding rudder surface and the fixed rudder surface, and after the pin puller is unlocked, the folding rudder surface is rotated to be unfolded in place under the action of the tension spring and locked by a in-position lock arranged in the fixed rudder surface. The longitudinal unfolding mechanism of the missile folding rudder surface achieves the rudder surface folding and unfolding function of a large-aspect-ratio rudder surface and small-diameter projectile.

The invention relates to a preparation method of an external thermal insulation composite material, comprising the following steps: (1) performing a hydrophobic treatment on a hydrophilic first reinforcement to obtain a hydrophobic first reinforcement; (2) connecting the hydrophobic first reinforcement and a hydrophilic second reinforcement to obtain a preform; and (3) impregnating the part of thepreform including at least the second reinforcement with a hydrophilic impregnant used for impregnation, and then drying to obtain the external thermal insulation composite material. The invention also relates to an external thermal insulation composite material prepared by the above preparation method. The an anti-scouring aerogel composite material prepared from the external thermal insulationcomposite material can be used for profile machining to obtain a component of a desired profile, or the overall external profile machining is carried out after assembling of an aircraft, so as to accurately ensure aerodynamic shape of the aircraft.

The invention relates to an unmanned helicopter complex-shaped composite-material integral oil tank structure. The unmanned helicopter complex-shaped composite-material integral oil tank structure comprises an oil tank cover, an oil tank body, a copper network, a left covering cap, a right covering cap, an oil level measurer, an oil filler, an oil outlet, an oil supply port and an oil return port.The oil tank body is divided into a middle section, a side face, a bottom and a front skirt, a rear skirt, a left skirt and a right skirt. An oil tank is placed below a rotor wing and on the periphery of a propeller hub, the oil tank cover and the side face are upper-side skins of a middle fuselage of a helicopter, the bottom of the oil tank body is placed on a truss in the middle fuselage, and the oil tank body is in bolted connection with a lower-side skin of the middle fuselage and an internal truss through the left skirt and the right skirt, in bolted connection with skins of a front fuselage and a rear fuselage through the front skirt and the rear skirt, in blind hole flange bolted connection with the oil level measurer through the bottom, the oil outlet, the oil supply port and theoil return port, in glue-joint connection with the oil tank cover through the middle section and the side face, and in indirect connection with the left covering cap and the right covering cap throughthe side face. The oil tank cover is in glue-joint connection with the oil filler, and the copper network is connected to the oil tank in a glued mode. The unmanned helicopter complex-shaped composite-material integral oil tank structure has the advantages of being high in strength, high in rigidity, low in weight, good in rectifying effect and capable of protecting a propeller hub main shaft andthe like.

The invention discloses a helicopter rotor blade based on a seamless trailing edge flap mechanism. The helicopter rotor blade comprises a blade body and a plurality of trailing edge flap modules, wherein the trailing edge flap module comprises a shell, a driver, a push-pull rod, a driving rod, a mounting platform, a first flexible hinge, a second flexible hinge, a trailing edge flap and a sealingpiece; the driver generates dynamic linear displacement output; the driver, the push-pull rod, the driving rod and the trailing edge flap are connected in sequence, and the force and displacement generated by the driver are output and transmitted to the trailing edge flap; and the trailing edge flap is connected with the mounting platform through the second flexible hinge, and the linear reciprocating motion of the push-pull rod is converted into the deflection motion of the trailing edge flap through the second flexible hinge. The helicopter rotor blade is applied to the trailing edge flaps of the helicopter to actively control a rotor wing, the aerodynamic load distribution of the blade is dynamically changed through dynamic deflection of the trailing edge flaps, the vibration load of the rotor wing of the helicopter is inhibited, or the noise of the rotor wing is inhibited through a sound field cancellation method.

The invention relates to the technical field of unmanned aerial vehicles, and provides a wing leading edge of a multifunctional solar aircraft. The wing leading edge comprises a leading edge body, energy storage batteries and an outer shell, wherein the leading edge body is made of foam materials, a battery cabin is formed in the leading edge body, and the energy storage batteries are placed in the battery cabin; the leading edge body is wrapped with the outer shell. According to the wing leading edge, the energy storage batteries can be kept warm, and the structural weight of the aircraft is reduced.

The invention relates to the field of unmanned aerial vehicles, in particular to a leading edge load mounting system of a flying wing stealth unmanned aerial vehicle. The leading edge load mounting system aims at solving the problem that a stealth aircraft cannot reflect radar waves because of inconvenient load mounting. The leading edge load mounting system of the flying wing stealth unmanned aerial vehicle comprises a concave mounting notch formed in the leading edge of a wing, a front spar, a leading edge wedge and a leading edge load hatch cover. The front spar is installed at the bottom of the mounting notch. The leading edge wedge is arranged on the front spar and covers the front spar. The leading edge load hatch cover covers the leading edge wedge and is in smooth transition with the body of the unmanned aerial vehicle. A mounting hole for mounting a load is formed in the leading edge wedge which is made of a conductive material and is in a convex wedge shape. A V-shaped cavity is formed between the leading edge wedge and the leading edge load cover. The leading edge load hatch cover is made of a wave-transmitting material. A load system has the capacity of carrying a built-in reflector, and can simulate all imaginary targets without affecting aerodynamic configuration.

The invention relates to the technical field of tactical missile shell section structures, in particular to a stringer housing which comprises a housing and a stringer, wherein the housing is rotary and comprises a covering skin and end frames located at both ends of the covering skin and integrally formed together with the covering skin; the stringer is connected with an inner wall of the housing in an axial direction of the housing. The housing of the invention is a rotary structure; the end frames and the covering skin form an integrated structure which has no projecting structure, can be completed by lathing, greatly shortens the processing cycle, and greatly improves the processing efficiency; positions of an upper end frame and a lower end frame are designed according to the size of an interface; the positions of the end frame and the covering skin adopt smooth transition, can be completed by lathing, greatly improve the processing efficiency, and reduce stress concentration. The stringer housing provided by the invention reduces the assembling difficulty and strength, effectively improves the assembling efficiency, also effectively reduces the weight of the missile structure, and improves the missile performance; the housing has high form-position tolerance grade, ensures aerodynamic shape of the missile, has attractive appearance, has no obvious splicing joints, and reduces the probability of appearance of rework events caused by appearance defects.

The invention discloses a combined type wind wheel blade and a wind generating set comprising the same. The combined type wind wheel blade is characterized in that docking end surfaces of adjacent combining sections are respectively connected with inner and outer flange plates, the inner and outer flange plates are inserted to combined, and side walls of inserting combined parts are joined so as to form a rectangular groove; the rectangular groove is divided into multiple sections, and the outer flange plate arranged between the adjacent sections is provided with an opening (wedge-shaped stripe inserting hole); the wheel blade also comprises multiple groups of wedge-shaped strip belts the number of which is equal to that of sections of the rectangular grooves, left and right wedge-shaped stripes of each group of wedge-shaped stripes are respectively inserted into the rectangular grooves from openings of the two ends of the rectangular groove of the same section in the reverse direction, so as to form interference fit with the rectangular grooves, and opening parts of subsections of the rectangular grooves are also fixedly provided with connecting end cap. The combined type wind wheel blade and the wind generating set provided by the invention have the advantages that the structure is simple, the connection is reliable, the assembly is simple and convenient, the production efficiency is high, the pneumatic appearance of the blade is guaranteed, the manufacture and transportation difficulty of a large-scale wind wheel blade are greatly reduced, the production and transportation costs are reduced, the production and transport efficiencies are improved, and local rapid assembly is realized.

A three-dimensional negative-stiffness elastic frame belongs to the technical field of aircraft deformation, solves the problem of high energy consumption of an existing bullet controlled by a deflection head, and comprises a cylinder wall provided with a plurality of hollowed-out parts, the hollowed-out parts are arranged in a delta shape, and the left-right direction of the delta shape is arranged in the circumferential direction of the cylinder wall; the cylinder can be a cylinder, an elliptical cylinder, a square cylinder, a polygonal cylinder and the like, or a straight cylinder, a bent cylinder, a reducing cylinder and the like in any shape, and can be arranged according to the requirements of the head of the aircraft; due to the fact that the hollowed-out parts arranged in the shape like the Chinese character'pin 'are formed in the cylinder wall and are arranged in the circumferential direction of the cylinder wall, the cylinder wall can have the axial elastic deformation function like a spring, and overall deflection can be achieved. The above technology is the prior art. The cylinder wall reserved around each hollowed-out part arranged in the shape like the Chinese character'pin 'comprises a top beam, a bottom beam and vertical beams on the two sides; the top beam and / or the bottom beam reserved around each hollowed-out part arranged in the shape like the Chinese character'pin' are outwards convex or inwards concave curved beams; the method is used for aircraft deformation.

The invention discloses an airship with an inflating skeleton expanding in air, which comprises an airship body 1, a tail wing 2, a tail thruster 3, a connecting element 7, a payload bay 8, a solar cell panel 9, a nose cone 10, a first auxiliary rope 11, an automatic inflating device 12, a second auxiliary rope 13 and two high-altitude balloons 14. The airship is characterized in that a dimensional element of the airship body 1 is skin 4, a longitudinal element is a rigid skeleton 5, and an annular element is an inflating skeleton 6; the inflating skeleton 6 is connected with the rigid skeleton 5 by the connecting element; and the skin 4 is fixedly connected on the outer surface of a grid formed by the rigid skeleton 5 and the inflating skeleton 6. The invention overcomes the defect that a soft bag in a folding state and a connecting part of a pod and other rigid structures are easily torn by strong wind and vortex in the lifting process.

The invention discloses a bullet with a helical deceleration device, comprising: a tail wing seat mounted on the tail of a casing, which evenly defines a plurality of grooves along the circumferential direction; a helical fin, comprising a fin and a fin seat arranged at the end of the fin. The pivot seat has gear teeth, which are accommodated in the groove; the fuze is installed in the hollow hole inside the tail wing seat, and the medicine box is accommodated therein; Racks are opened on the spaced prism surfaces, and the edge surfaces of the racks correspond one-to-one with a plurality of grooves on the tail wing seat; after the fuze is ignited, the charge box is detonated, and a force is generated to push the transmission disc to move axially, so as to pass through its outer surface. The rack drives the pivot seat meshed with it to rotate, so that the helical fins are spread by covering the surface contour of the bullet shell, and the helical deceleration of the bullet can be realized under the action of aerodynamic force. The invention adopts the helical fins to realize the effective and stable deceleration of the bullet, and solves the problems of complex structure and low reliability of the existing bullet deceleration device.

The invention discloses a combination type wind turbine blade and a wind generating set comprising the same. The butt end surfaces of two adjacent combination sections of the blade are connected with an inner flange plate and an outer flange plate respectively; the inner flange plate and the outer flange plate are in insertion fit with each other, and the side walls of insertion fit parts are intersected to form a plurality of rectangular grooves; each rectangular groove is divided into an upper section and a lower section through a front edge opening and a rear edge opening in the outer flange plate; the wind turbine blade further comprises a plurality of groups of wedge-shaped strips; a left wedge-shaped strip and a right wedge-shaped strip in each group of wedge-shaped strips are reversely inserted into the corresponding rectangular groove from openings in two ends of the same rectangular groove respectively and are in interference fit with the corresponding rectangular groove; positioning plugs are further fixedly arranged at the front edge opening and the rear edge opening. The combination type wind turbine blade is simple in structure, high in connection strength, high in reliability, easy, convenient and quick to mount, good in torsion resistance after connection, and good in pneumatic appearance; the manufacturing difficulty and the transporting difficulty of large-sized wind turbine blades can be reduced, the production cost and the transporting cost are reduced, the production efficiency and the transporting efficiency are improved, and on-site quick assembly is realized.

The invention relates to the technical field of aircraft design, in particular to a ram air entrainment structure. The structure comprises an air entrainment device, a plurality of air entrainment holes are arranged along the axial direction, and one end facing the outside of the machine is conformal with the fuselage; One end of the supporting device is fixedly connected with one end of the air guide structure facing the inside of the machine, and is used for supporting the air guide structure and communicating the air guide hole with the inside of the machine body. The structure is arrangedin an air inlet opening of the airframe, no protrusion is formed on the exterior of the airframe after installation, and the structure is highly integrated with the airframe so as to maintain the aerodynamic shape of the airplane, thereby reducing the flight resistance of the airplane, stabilizing the flight of the airplane, and avoiding damaging the stealth performance of the airplane to a certain extent; In addition, the structure is simple, easy to manufacture, easy to install, easy to replace the existing ramp air inlet structure design.

The invention relates to the technical field of tactical missile shell section structures, in particular to a stringer housing which comprises a housing and a stringer, wherein the housing is rotary and comprises a covering skin and end frames located at both ends of the covering skin and integrally formed together with the covering skin; the stringer is connected with an inner wall of the housing in an axial direction of the housing. The housing of the invention is a rotary structure; the end frames and the covering skin form an integrated structure which has no projecting structure, can be completed by lathing, greatly shortens the processing cycle, and greatly improves the processing efficiency; positions of an upper end frame and a lower end frame are designed according to the size of an interface; the positions of the end frame and the covering skin adopt smooth transition, can be completed by lathing, greatly improve the processing efficiency, and reduce stress concentration. The stringer housing provided by the invention reduces the assembling difficulty and strength, effectively improves the assembling efficiency, also effectively reduces the weight of the missile structure, and improves the missile performance; the housing has high form-position tolerance grade, ensures aerodynamic shape of the missile, has attractive appearance, has no obvious splicing joints, and reduces the probability of appearance of rework events caused by appearance defects.

The invention discloses a differential motion operation mechanism for follow-up spoilers of an airplane. According to the differential motion operation mechanism, the lateral controllability of the airplane at a high attack angle is improved, and the aileron effect is increased. The differential motion operation mechanism adopts a follow-up operation manner and can automatically stretch or retractalong the motions of ailerons, when the ailerons lean downwards from a neutral position, the spoilers are still adhered to the surface of wings, and the aerodynamic configurations of the surfaces ofthe wings are maintained, so that the occupied space of the wings is small, and a system is convenient to configure and easy to mount and debug.

The invention discloses a high-rigidity radome structure for counteracting an aerodynamic suction effect. The radome structure comprises a skin, a perforated honeycomb, an antenna and a U-shaped sealing element; the skin is arranged outside the antenna, and a vacuum cavity is formed between the skin and the antenna; the perforated honeycomb is arranged in the vacuum cavity; two symmetrical U-shaped sealing pieces are arranged on the edge of the root of the perforated honeycomb; openings of the two U-shaped sealing pieces are connected; the vacuum cavity is completely sealed; a gas valve is mounted on each U-shaped sealing piece; a thickening part is arranged in the root area of the skin; the outer skin of a radome is supported by a honeycomb; when the radome is subjected to pneumatic pressure; the radome skin and the filling honeycomb are supported by an internal antenna, and thus the aerodynamic shape of the radome is maintained and the electromagnetic wave transmittance of the radomeis improved; through honeycomb internal vacuumizing and preset inward suction force, when the outer surface of the radome is subjected to pneumatic suction force, internal vacuum counteracts part ofpneumatic suction force outside the radome, the bearing capacity of the radome is improved, and deformation of the radome is reduced.

The invention relates to a forming method of a railway vehicle cowl prepared from a carbon fiber reinforced resin matrix composite material. The forming method comprises the step of forming the vehicle cowl by using a cowl assembling die and specifically comprises the following steps of performing laying and curing forming by adopting a carbon fiber braid; assembling a metal embedded part and carbon fiber composite material reinforcing bars on a cured cowl covering in the cowl assembling die; and then demolding. According to the preparation method disclosed by the invention, by optimizing and innovating the forming die structural design, the laying structural design, the forming process design and the resin formula system design of the railway vehicle cowl, the prepared cowl has excellent integrated properties and has the advantages of low porosity, high specific strength, high specific rigidity, fatigue resistance, good performance stability, light weight and the like; and the integration of the structure and functions is realized, and use requirements of high-speed railway vehicles are met.

The invention belongs to the technical field of airplane wind tunnel tests, and discloses a pneumatic test device based on large-size control surface hinge moment, which comprises a vertical fin, a horizontal tail connecting plate, a horizontal tail connecting seat, a wind tunnel connecting seat, a horizontal tail stabilizer, spherical hinges and an elevator. The vertical fin is fixedly connected with the horizontal tail stabilizer through the horizontal tail connecting plate and the horizontal tail connecting seat, the horizontal tail stabilizer is connected with the elevator through the at least two spherical hinges, and the bottom of the elevator is connected. The whole device is installed in a wind tunnel through the wind tunnel connecting seat. According to the pneumatic test device based on the large-size control surface hinge moment, through the design of the spherical hinges, the technical problem of connection of an elevator and a stabilizer in a large-size control surface hinge moment test is effectively solved, and meanwhile the measurement precision of a balance is guaranteed.

The invention belongs to the technical field of mechanical processing and manufacturing, and discloses a method for manufacturing a super-long thin-walled aluminum alloy sandwich grid rib double-skin integral tail, comprising: adopting ZL205AA high-strength aluminum alloy to form the integral tail; The integral tail is heat treated; the heat treated integral tail is processed by numerical control. The manufacturing method of the super-long thin-walled aluminum alloy sandwich grid rib double-skin integral empennage provided by the present invention can improve the overall quality of precision molding and manufacturing, reduce the manufacturing cost and shorten the manufacturing cycle.

The invention discloses a miniaturized radar fuze structure which consists of a transmitting antenna, a millimeter-wave transceiving module, a video amplifier, a digital signal processor and a receiving antenna. The millimeter-wave transceiving module, the video amplifier and the digital signal processor are cylindrical, and are sequentially connected and fastened two by two in the peripheral direction; the transmitting and receiving antennas are fastened to the outer surfaces of the millimeter-wave transceiving module, the video amplifier and the digital signal processor, and are in cross-shaped symmetry arrangement. According to the structure, not only is the aerodynamic shape of a cabin is guaranteed, but also the utility rate of space is improved greatly, the overall structure is compact, and the utility rate of space is high. The fuze structure is small, light and modularized, and complies with the trend in the development of modern weapon systems.

The invention discloses a wing folding mechanism that can be folded at a large angle with double rotation driven by a straight line, comprising a connecting shaft, an inner and outer wing connecting joint installed on the connecting shaft, and a telescopic folding mechanism connected with the connecting shaft; the telescopic folding mechanism includes Two telescopic folding devices with the same structure and hinged with the connecting shaft; one end of the telescopic folding device away from the connecting shaft is hinged to the outer wing, and the other end of the telescopic folding device away from the connecting shaft is hinged to the inner wing. The connecting joints of the inner and outer wings are respectively hinged with the inner wing lugs on the inner wing and the outer wing lugs on the outer wing; a gap for installing the connecting shaft and the telescopic folding mechanism is formed between the inner wing lugs and the outer wing lugs. Chamber. The beneficial effects of the present invention are: the mechanism of the present invention can ensure the aerodynamic shape of the wing itself without affecting the flight performance of the aircraft, and the folding angle is large, the height after folding is low, and the occupied space is small; the present invention is operable, simple in principle, and applicable Wide range of features, short design cycle and low difficulty.

The invention relates to a method for preparing an anti-scouring aerogel composite material. The method comprises: (1) connecting a hydrophobic first reinforcement to a hydrophilic second reinforcement to obtain a preform, (2) dipping the preform part at least comprising the second reinforcement in a hydrophilic first impregnating material subjected to densification treatment and drying the preform to obtain a composite material, and (3) carrying out hydrophilic treatment on the preform part at least comprising the hydrophobic first reinforcement to obtain the anti-scouring aerogel composite material. The invention also relates to the anti-scouring aerogel composite material obtained by the method. The anti-scouring aerogel composite material is suitable for profile machining so that a member with a desired profile is obtained, or the whole outer profile is subjected to machining after aircraft assembling so that the aerodynamic shape of the aircraft is accurately ensured.

The invention discloses a sectional type fan blade connecting structure, a sectional type fan blade and an assembling method. The sectional type fan blade connecting structure is characterized in that a first socket piece, a second socket piece and a fixing assembly are arranged, and the first socket piece is arranged at one end of a first blade section; the second socket component is arranged at one end of the second blade section; the first socket component and the second socket component are opposite and are arranged in a mutually crossed and staggered manner to form a connecting part, and the fixing component is arranged at the connecting part to fix the first socket component and the second socket component. According to the connecting structure, connection of the segmented blades is more reliable, the additional mass of the connecting area is light, and factory production and wind field on-site operation are facilitated.

The invention relates to the technical field of aircrafts, in particular to a honeycomb supporting structure comprising a flexible skin and a preparation method of the flexible skin, the honeycomb supporting structure is a non-full-thickness penetrating type structure and comprises a zero-Poisson-ratio honeycomb, a web, a center side strip, a rear edge section, a front edge mounting section, an inclined side pull rod and the flexible skin; the flexible skin is formed by compounding a one-way elastic nylon fabric and silicone rubber; the zero-Poisson-ratio honeycombs are arranged on the upper surface and the lower surface of the honeycomb supporting structure; the webs are arranged between the zero-Poisson-ratio honeycombs on the upper surface and the lower surface in the spanwise direction, and a through groove is formed in the middle of each web; a supporting plate is arranged at the front end of the rear edge section, a sliding groove in the spanwise direction is formed in the middle of the supporting plate, and connecting pieces are arranged on the two sides of the supporting plate respectively. The through groove and the sliding groove are connected through a driving shaft, and the honeycomb supporting structure is driven to be bent and subjected to torsional deformation. According to the invention, the problems of poor pneumatic effect and low tear resistance can be effectively solved.